Fuel injector for a gas turbine engine

ABSTRACT

The invention comprises a gas turbine engine having a fuel injector comprising a main body having therein a plurality of ducts terminating in respective nozzles, fuel supply means for each of said nozzles, a hollow annular cowl mounted in a radially spaced relationship about the said main body so as to define therebetween an air inlet, baffle means disposed downstream of said nozzles, and valve means associated with a said duct for selectively connecting the latter to its respective fuel supply means or to a source of high pressure air, so that, in operation of the injector, the fuel, or fuel and air, emitted by said nozzles passes between said cowl and said baffle means, and simultaneously therewith a flow of high pressure air passes from the compressor stage of the gas turbine engine through said air inlet.

United States Patent Carlisle 1 1 Sept. 12, 1972 [22] Filed:

[54] FUEL INJECTOR FOR A GAS TURBINE ENGINE [73] Assignee: Rolls-Royce Limited,

Derbyshire, England Dec. 4, 1970 [21] Appl. N0.: 95,124

Derby,

[30] Foreign Application Priority Data Dec. 9, 1969 Great Britain ..60,035/69 [52] US. Cl ..60/39.74 R, 60/3946, 239/400, 239/407 [51] Int. Cl. ..F02c 3/22 [58] Field of Search ..60/39.74 R, 39.27, 39.46; 239/428, 400, 407

[56] References Cited UNITED STATES PATENTS 2,595,759 5/1952 Buckland ..60/39.74 R 2,635,425 4/ 1953 Thorpe ..60/39.74 R

Cummings ..60/39.74 R Carlisle ..60/39.74 R

[57] ABSTRACT The invention comprises a gas turbine engine having a fuel injector comprising a main body having therein a plurality of ducts tenninating in respective nozzles, fuel supply means for each of said nozzles, a hollow annular cowl mounted in a radially spaced relationship about the said main body so as to define therebetween an air inlet, baffle means disposed downstream of said nozzles, and valve means associated with a said duct for selectively connecting the latter to its respective fuel supply means or to a source of high pressure air, so that, in operation of the injector, the fuel, or fuel and air, emitted by said nozzles passes between said cowl and said baffle means, and simultaneously therewith a flow of high pressure air passes from the compressor stage of the gas turbine engine through said air inlet.

9 Claims, 3 Drawing Figures PATENTED 3,690,093

SHEET 2 OF 2 P I Inventor BEMS mam-(D QRRLASLE (210104014, 1 uhwfitlorney FUEL INJECTOR FOR A GAS TURBINE ENGINE This invention is concerned with a fuel injector for a gas turbine engine.

According to the present invention, there is provided a gas turbine engine having a fuel injector comprising a main body having therein a plurality of ducts terminating in respective nozzles, fuel supply means for each of said nozzles, a hollow annular cowl mounted in a radially spaced relationship about the said main body so as to define therebetween an air inlet, a baffle means disposed downstream of said nozzles, and valve means associated with a said duct for selectively connecting the latter to its respective fuel supply means or to a source of high pressure air, so that, in operation of the injector, the fuel, or fuel and air, emitted by said nozzles passes between said cowl and said baffle means, and simultaneously therewith a flow of high pressure air passes from the compressor stage of the gas turbine engine through said air inlet.

The fuel injector is of the kind which is the subject of British Pat. application No. 59965/ 69.

It will be appreciated that in a gas turbine engine having a fuel injector according to the present invention, the risk that a nozzle which is connected to the duct associated with said valve means and which is not in use at any particular time will carbon up is reduced by reason of the fact that at such time a flow of air through said duct will prevent entry thereinto of fuel from the operative nozzle or nozzles.

Preferably the said nozzles are concentric and the outermost nozzle is a gaseous fuel nozzle, the other nozzle or nozzles being liquid fuel nozzles.

In a preferred embodiment, said valve means is associated with said gaseous fuel nozzle only.

The fuel supply means are preferably controlled so as to permit simultaneous supply of gaseous fuel to said outermost nozzle and liquid fuel to the other nozzle(s).

There may be, in said central body, a relatively small diameter central liquid fuel duct terminating in a nozzle for use in pilot fuel combustion which is concentrically surrounded by a relatively large cross-sectional area annular liquid fuel duct terminating in a nozzle for use in main fuel combustion.

The baffle means is preferably supported by said cowl and comprises a member having a substantially conical surface the axis of which is common with the axis of said nozzles and the apex of which is disposed towards said nozzles.

Preferably said member has a plate secured to its downstream end which plate serves as a heat sink for the prevention of carbon deposits on said member.

Optionally, the downstream end of the cowl is divergent in a downstream direction.

Advantageously, at least a portion of said cowl is convergent in a downstream direction.

Preferably, the downstream end of said cowl is apertured to permit a flow of high pressure air to pass therethrough and into the air inlet.

The invention is illustrated, merely by way of example, in the accompanying drawings, in which:

FIG. 1 is a diagrammatic view, partly in broken-away longitudinal section, of a gas turbine engine provided with a fuel injector according to the present invention,

FIG. 2 is a broken away longitudinal sectional view on a larger scale of the fuel injector shown in FIG. 1, and

FIG. 3 is a section taken on the line 33 of FIG; 2.

Referring to the drawings, a gas turbine engine 10 comprises in flow series one or more compressors 11, combustion equipment 12, and one or more turbines 13, the turbine exhaust gases being directed to atmosphere through an exhaust duct 14 which may terminate in a jet nozzle.

The combustion equipment 12 comprises a plurality of circumferentially spaced apart flame tubes 15. Each of the flame tubes 15 is provided at its upstream end with a fuel injector 16.

Each fuel injector 16 (FIG. 2) comprises a central body 17 having secured thereto a nut 18 the downstream end of which is shaped to form a relatively small, centrally disposed nozzle 20 for the so-called pilot fuel, the nozzle 20 being concentrically surrounded by a relatively large annular nozzle 21 formed by the nut 18 for the so-called main fuel, the pilot and main fuels in this preferred embodiment being liquid fuels. The central body 17 is secured centrally in one arm of a substantially L-shaped fuel feed arm 19 in which are located three fuel supply ducts 43, 44 and 45. The left hand duct 43 (as seen in FIG. 2) communicates with a pilot fuel passage 24 in the central nozzle 20 and the central fuel supply duct 44 communicates with the main fuel passage 28 in the large annular nozzle 21. The right hand duct 45 communicates with an annular passage 35 formed between the central body 17 and a shroud 22a brazed to the limb of the fuel feed arm 19 which is substantially co-axial with the passages 28, 35 to form a nozzle 22 therewith. The central body 17 and the fuel feed arm 19 together form said main body.

The downstream end of thenozzle 20 is provided with circumferentially spaced drillings 23 which communicate at their upstream ends with a pilot fuel passage 24 in the body 17 and which communicate at their downstream ends with the interior of a hollow annular cowl 25. The shroud 22a supports the cowl 25 by means of dowels 29 so that the latter is radially spaced from the nozzle 22. The cowl 25 carries three equi-angularly spaced supporting members 26. Each member is substantially Y-shaped in section and is secured, e.g., by brazing, to a substantially conical baffle member 27 to support the latter. The cowl 25' has an open upstream portion which defines, with the external surface 7 of the shroud 22a an air inlet 30. The cowl 25 converges from its upstream end 30 to an intermediate point 31 which is substantially co-planar with the downstream end of the nozzle 20; the cowl 25 then diverges from the intermediate point 31 to an open downstream end 32 which is provided with throughgoing apertures 33 which communicate with the flow passage 35 defined between the central body 17 and nut 18 on the one hand and the shroud 22a and the cowl 25 on the other hand.

The baffle member 27 is disposed inwardly of the downsneam end 32 of the cowl 25, and the axis of its conical surface substantially coincides with the axis of the nozzles 20, 21, 22, with the apex of the cone being disposed towards the nozzle 20.

The downstream end of the baffle member 27 has a metallic heat shield member 36 secured thereto. It will be appreciated that this member 36 is in operation exposed to the internal temperature of the flame tube and will thus transmit heat to the conical baffle member 27 to a sufficient extent to prevent any accumulation of carbon deposits thereon.

Air which has been compressed by the compressor or compressors 11 is in operation forced through the inlet in the cowl 25.

The passage of either gas or liquid fuel takes place through the cowl 25 simultaneously with a flow of air, the flow of air through the cowl 25 not only effecting atomization of the fuel but also helping to ensure that the fuel is burned away from the nozzles 20, 21, 22. Thus carboning up of any of the nozzles 20, 21, 22 which happens to be out of action at any particular time is reduced. The baffle member 27 assists in atomizing the fuel, the fuel and air being directed into a desired direction between the baffle member 27 and the downstream end 32 of the cowl 25.

The gas fuel passage 45 communicates at one end with the flow passage 35, and at the other end with a gas manifold 52 (FIG. 1). The liquid fuel ducts 43 and 44 respectively communicate with the pilot fuel passage 24, via an annular chamber 55, and with the main fuel passage 28. Both the liquid ducts 43, 44, communicate with a liquid fuel manifold 56 (FIG. 1) so as to receive fuel therefrom.

The manifolds 52, 56 are connected by respective lines 57, 58 to a dual fuel flow control unit 60 which has means (not shown) for enabling gaseous fuel and liquid fuel to be used either simultaneously or separately from each other.

It has been found, however, that when the fuel injector 16 is operated on liquid fuel alone, the problem arose that liquid fuel sometimes splashed back into the annular passage and hence to the gas passage 45 wherein it spontaneously ignited. This spontaneous ignition can, of course, damage the injector structure and has to be prevented. Furthermore, in operation, the prevailing pressures in the various combustion chambers of a given engine are not always equal, and this gives rise to the possibility of hot combustion products flowing back into the annular passage 35 of an injector of a relatively high pressure combustion chamber and from there, via the gas fuel manifold 52 into a relatively low pressure combustion chamber. This is clearly undesirable.

These problems are overcome according to the illustrated embodiment of this invention by the provision of a pair of shut-off valves 61, 62 (FIG. 1) connected to the gas fuel line 57. The valves 61, 62 are so arranged that when one is open, the other one is shut and vice versa. The valve 61 is disposed in a line 63 which connects the gas fuel line 57 with a source of compressed air which in the illustrated embodiment is tapoff air from the outlet of the compressor 11. The valve 62 is upstream of the connection between the lines 57 and 63, and is disposed in the gas fuel line 57.

In operation, when liquid fuel only is being burnt, the gas fuel supply is shut off by means of the valve 62, while the valve 61 is opened to supply a continuous flow of compressed air from the line 63 to the gaseous 4 fuel nozzle 22 by way of the line 57, the gas manifold 52 and the passage 45. t

I claim:

1. A gas turbine engine having a compressor stage and a fuel injector comprising a main body having therein a lurality of ducts terminating in at least one liquid fue nozzle and at least one gaseous fuel nozzle,

liquid fuel supply means for the liquid fuel nozzle and gaseous fuel supply means for the gaseous fuel nozzle, a hollow annular cowl mounted in radially spaced relationship about said main body, said cowl and said main body being spaced to define an air inlet, bafile means disposed downstream of said nozzles, valve means associated with the duct which terminates in the gaseous fuel nozzle, a source of high pressure air, means for selectively connecting the gaseous fuel nozzle duct to the gaseous fuel supply means or to said source of high pressure air so that the liquid and high pressure air can flow from the liquid fuel nozzle and the gaseous fuel nozzle respectively between said cowl and said baffle means together with a flow of high pressure air from said compressor stage of said engine which flows through said air inlet.

2. A gas turbine engine as claimed in claim 1 in which the said nozzles are concentric and the outermost nozzle is the gaseous fuel nozzle.

3. A gas turbine engine as claimed in claim 2 wherein the fuel supply means are controlled so as to permit simultaneous supply of gaseous fuel to said outermost nozzle and liquid fuel to at least the one liquid fuel nozzle.

4. A gas turbine engine as claimed in claim 1 in which in said main body there is a relatively small diameter central liquid fuel duct terminating in a nozzle for use in pilot fuel combustion which is concentrically surrounded by a relativelylarge cross section annular liquid fuel duct terminating in a nozzle for use in main fuel combustion.

5. A gas turbine engine as claimed in claim 1 wherein the baffle means is supported within said cowl and comprises a member having a substantially conical surface the axis of which is common with the axis of said nozzles and the apex of which is disposed towards said nozzles.

6. A gas turbine engine as claimed in claim 5 wherein the said member has a plate secured to its downstream end which plate serves as a heat sink for the prevention of carbon deposits on said member.

7. A gas turbine engine as claimed in claim 1 wherein the downstream end of the cowl is divergent in a downstream direction.

8. A gas turbine engine as claimed in claim 1 wherein at least a portion of said cowl is convergent in a downstream direction.

9. A gas turbine engine as claimed in claim 1 wherein the downstream end of said cowl is formed with an aperture to permit a flow of high pressure air to pass therethrough and into a flow passage between said cowl and said main body. 

1. A gas turbine engine having a compressor stage and a fuel injector comprising a main body having therein a plurality of ducts terminating in at least one liquid fuel nozzle and at least one gaseous fuel nozzle, liquid fuel supply means for the liquid fuel nozzle and gaseous fuel supply means for the gaseous fuel nozzle, a hollow annular cowl mounted in radially spaced relationship about said main body, said cowl and said main body being spaced to define an air inlet, baffle means disposed downstream of said nozzles, valve means associated with the duct which terminates in the gaseous fuel nozzle, a source of high pressure air, means for selectively connecting the gaseous fuel nozzle duct to the gaseous fuel supply means or to said source of high pressure air so that the liquid and high pressure air can flow from the liquid fuel nozzle and the gaseous fuel nozzle respectively between said cowl and said baffle means together with a flow of high pressure air from said compressor stage of said engine which flows through said air inlet.
 2. A gas turbine engine as claimed in claim 1 in which the said nozzles are concentric and the outermost nozzle is the gaseous fuel nozzle.
 3. A gas turbine engine as claimed in claim 2 wherein the fuel supply means are controlled so as to permit simultaneous supply of gaseous fuel to said outermost nozzle and liquid fuel to at least the one liquid fuel nozzle.
 4. A gas turbine engine as claimed in claim 1 in which in said main body there is a relatively small diameter central liquid fuel duct terminating in a nozzle for use in pilot fuel combustion which is concentrically surrounded by a relatively large cross section annular liquid fuel duct terminating in a nozzle for use in main fuel combustion.
 5. A gas turbine engine as claimed in claim 1 wherein the baffle means is supported within said cowl and comprises a member having a substantially conical surface the axis of which is common with the axis of said nozzles and the apex of which is disposed towards said nozzles.
 6. A gas turbine engine as claimed in claim 5 wherein the said member has a plate secured to its downstream end which plate serves as a heat sink for the prevention of carbon deposits on said member.
 7. A gas turbine engine as claimed in claim 1 wherein the downstream end of the cowl is divergent in a downstream direction.
 8. A gas turbine engine as claimed in claim 1 wherein at least a portion of said cowl is convergent in a downstream direction.
 9. A gas turbine engine as claimed in claim 1 wherein the downstream end of said cowl is formed with an aperture to permit a flow of high pressure air to pass therethrough and into a flow passage between said cowl and said main body. 